The invention relates to a rocket engine with a central body enclosing a combustion chamber, a nozzle and an expansion chamber mounted in a support structure.
In a known rocket engine as shown in FIG. 1, a converging combustion chamber 1 and a nozzle 2 and a diverging expansion chamber 4 are positioned between a fuel injection head 3 and an expansion chamber extension 4xe2x80x2. The combustion chamber 1 and the nozzle 2 are the thermo-mechanically most heavily loaded components of such a rocket engine. The fuel injection head 3 injects fuels, for example hydrogen (H2) and oxygen (O2) or kerosene and nitric acid (MMH) and an oxidizer into the combustion chamber 1 for combustion. Combustion gases generated in the combustion chamber 1 converge into the nozzle 2 at very high temperatures and under very high pressures. Additionally, substantial pressure surges occur well above the normal operating conditions during certain operating ranges such as at ignition and when the engine is turned off. In addition, there is a high gas velocity in the expansion chamber 4 and its extension 4xe2x80x2 diverging away from the nozzle 2. The combustion gases are expanded in the diverging section 4, 4xe2x80x2 to the nozzle ultimate pressure of the engine.
A refractory high alloy steel with a high mechanical strength and a high temperature resistance must be used for constructing a conventional combustion chamber for a rocket engine due to the high temperatures and high pressures mentioned above. These high alloy steels, for example Inconel(copyright), however still have the disadvantage of losing strength at temperatures of about 800xc2x0 C. and higher in such a way that additional active cooling of the engine is necessary. This cooling is provided by appropriate cooling channels 6 cut or machined into the wall 5 that encloses the combustion chamber 1, the nozzle 2 and the expansion chamber 3. After machining, the cooling channels must be closed again at the channel surface by chemically applied material so that a coolant can flow through closed coolant channels. Such a formation of the cooling channels requires a difficult and cost intensive production.
In view of the foregoing it is the aim of the invention to achieve the following objects singly or in combination:
to provide a power unit, more specifically a rocket engine body which encloses a converging combustion chamber, a nozzle and a diverging expansion chamber, preferably surrounded by an insulation, whereby such a body has a high temperature resistance, a high mechanical strength against pressure loads and a high abrasion resistance even at a low density, at a high heat conduction, and at a low thermal expansion;
to substantially reduce the weight of rocket engines, particularly the weight of the above mentioned body that forms the converging combustion chamber, the nozzle proper, and the diverging expansion chamber;
to use easily machinable or moldable materials for producing the critical rocket engine components to permit substantially unlimited geometric and shape variations; and
to use materials that facilitate any machining operations as compared to high alloy steels which are difficult to machine.
According to the invention the rocket body that forms the combustion chamber, the nozzle, and the expansion chamber, and preferably also the insulation surrounding the body, are made of a ceramic material, namely carbon fiber reinforced silicon carbide composite material referred to herein as C/SiC. The C/SiC composite material body may be formed of a plurality of body sections which are ceramically or mechanically connected with each other. Alternatively, the body is formed as a monolithic structure. In the embodiment with a plurality of body sections these sections are bonded to each other by a silicon bonding layer which is siliconized to form a bonding seam or seams. The insulation, if used may also be provided in sections or may be part of the monolithic rocket engine structure.
According to a preferred embodiment of the invention the body forming the combustion chamber, the nozzle, the expansion chamber and the insulation surrounding the body, are made of the same C/SiC composite material to thus provide the monolithic structure, whereby the temperature stability of the engine is increased while simultaneously reducing the weight of the engine with its insulation compared to conventional engines of comparable size made of high alloy steel.
It was found that C/SiC has excellent strength characteristics up to very high temperatures which allow using this material under difficult operating conditions. Additionally, C/SiC offers a low density in combination with a high resistance to abrasion, a high resistance to oxidation, an excellent temperature stability even at temperature fluctuations, and a high thermal fatigue strength combined with an absolute gas and liquid impermeability. A great variety of geometric configurations are easily realized with C/SiC as used according to the invention because machining is easy compared to high alloy steels. The present rocket engine component has an excellent thermal strength and a high heat conductivity that is even adjustable by selecting the porosity of the C/SiC material. This characteristic of a high heat conductivity allows for correspondingly low cooling capacities. Certain engines according to the invention can completely do without cooling due to the high thermal strength and excellent heat conductivity.
A distinction is made between C/SiC reinforced by long, continuous length carbon fibers, and C/SiC reinforced by short-length carbon fibers or whiskers. The material with continuous length carbon fibers can be laminated, pressed, or rolled. The resulting components have a particularly high mechanical strength combined with a particularly low density. The oxidation resistance of C/SiC can be improved by providing a suitable surface sealing. Such a sealing is not required for components made of short-fiber reinforced C/SiC because the material is especially resistant against oxidation and corrosion. Furthermore, C/SiC offers an extremely good heat dissipation due to its good heat conduction. This excellent characteristic is combined with an especially high thermal shock resistance. Moreover, C/SiC is particularly suitable for mechanical machining of blanks in a green state, whereby the combustion chamber cross-section, the nozzle cross-section and the expansion chamber cross-section can be machined to any geometry or configuration, either out of a monolithic blank or out of individual blank sections which can be formed into a lining for the combustion chamber, the nozzle, and the expansion chamber.
If the rocket engine body is formed of several sections, the body sections and the insulation can be interconnected by a silicon layer interposed between these sections to form a siliconized bonding between the sections, whereby the desired monolithic structure is obtained. This bonding of sections to each other either of the body and/or of the insulation is especially suitable for short-carbon fiber reinforced C/SiC, because the internal surface areas of the sections surrounding the combustion chamber, the nozzle, and the expansion chamber, are more easily machined prior to bonding the sections to each other. The siliconization or silicon infiltration for the bonding can then follow as a separate operation. In both embodiments, the monolithic structure as well as the structure with bonded seams can be easily connected by flanges to the fuel injection head and to the extension of the expansion chamber. The flanges are preferably also made of C/SiC. Further, mechanical machining can be used to easily form cooling channels with round, rectangular, or slot-shaped cross-sections in the body and/or in the insulation portion or layer.
According to a further embodiment of the invention the inner wall of the rocket engine body surrounding the combustion chamber, the nozzle and the expansion chamber is lined with C/SiC sections which are bonded to each other by siliconizing to form a monolithic body. Such a monolithic body protects a support structure of the body and of any heat insulation. The support structure is preferably made of metal. The support structure in turn protects the rocket engine body against pressure overloads. The protection against heat overload may be further improved by cooling channels through which fuels such as hydrogen may flow as a coolant. An insulation layer preferably also made of C/SiC or out of carbon fiber felts or of a graphite film or combinations of these materials further reduces the heat and pressure load on the metallic support structure. The insulation materials can be connected to the C/SiC body by intermediate spacers between the insulation and the body. The spacers are preferably also made of C/SiC in order to form the desired monolithic structure. The spacers may form flow passages for a coolant between the insulation and the rocket engine body.
The density and porosity of C/SiC materials can be advantageously adjusted while the siliconization is taking place by controlling the quantity of silicon or silicon carbide added for the bonding of the material components by siliconization. For example, C/SiC with a high density and a low porosity may be used as a thermo-mechanical structure or as a lining for the above mentioned rocket engine body, while a C/SiC material with a low density and a high porosity can be used as the thermal insulation, preferably in the form of a layer or layers or coatings.
The above described insulation of the support structure against heat emanating from the C/SiC rocket engine body is achieved either by cooling channels or by the above mentioned thermal insulation layers that may be formed of C/SiC material and/or carbon fiber felts and/or graphite films. The insulation may be used alone or in combination with the cooling channels. This thermal protection minimizes the heat and pressure loads on the metallic support structure. As a result, a reduced deformation of the rocket engine body, especially in the nozzle cross-section area is assured. Due to the resulting smaller loads on the metallic support structure, the latter can be substantially simplified, which in turn means a simplified production because the previously necessary and expensive material application by chemical action for closing up the machined cooling channels are no longer necessary.
The rocket engine body made of C/SiC that surrounds the combustion chamber, the nozzle and the expansion chamber is supported by the above-mentioned support structure, preferably made of metal, for example in the form of a housing which transfers any static and dynamic in-flight loads to the fuselage. The housing provides or supports the mechanical connections, for example by flanges, to the fuel injection head and to the expansion chamber extension. In addition, the support structure connects the cooling system of the fuel injection head with the cooling system of the expansion chamber through any cooling channels in the support structure and/or in the insulation and/or in the rocket engine body.
Due to the gas and liquid impermeability of C/SiC materials, open cooling channels can be machined into the metallic support structure or into the insulation or into the rocket engine body of C/SiC material. These initially open cooling channels are closed up when the C/SiC body is inserted into the support structure without any other efforts to close the cooling channels along their open sides. Depending on the type of system or engine used, the inner profile of the combustion chamber is either made of C/SiC section blanks which are siliconized to form a monolithic structure as mentioned. However, it is also possible to produce the rocket engine body from a single piece blank. Machining the section blanks is performed prior to bonding the sections together. The C/SiC sections can themselves, if necessary, be provided with cooling channels in order to still more efficiently dissipate heat. Any insulation interposed between the radially outwardly facing surfaces of the body of C/SiC material and the support structure must be connected to the support and the rocket engine body. Suitable connecting elements are used to mount the body of C/SiC material in the support structure, for example by the above mentioned flanges of the rocket engine body whereby the body flanges are secured to further flanges of the support structure.